Thermal analysis of a gas turbine cycle for a turbojet engine

International Journal of Advanced Research in Engineering and Technology (IJARET), ISSN 0976 – 6480(Print),
ISSN 0976 – 6499(Online) Volume 5, Issue 10, October (2014), pp. 21-33 © IAEME
21
THERMAL ANALYSIS OF A GAS TURBINE CYCLE FOR
A TURBOJET ENGINE
1
Rana Adil Abdul-Nabe, 2
Dr. Mohammad Tariq
1
Teacher, Ministry of Higher Education,
Institute of the Musaib Technical Institute, Republic of Iraq
2
Assistant Professor, Department of Mechanical Engineering,
SSET, SHIATS-DU, Allahabad, India, 211007
ABSTRACT
This paper deals with the application of a gas turbine engine for a single spool turbojet
engine. The design and operation theories of the components used in the jet engines are complicated.
The complexity of thermodynamic analysis makes it impossible to mathematically solve the
optimization equations involve in turbojet engine cycle. It was expected that these tools would help
in predicting the performance of individual components such as compressors, combustion chamber,
turbine, nozzle etc. The present work based on detailed parametric thermodynamic analysis of the
single spool turbojet engine with transpiration air cooling techniques for turbine blade cooling. The
analysis has been carried out by selecting models for different components of engines. Software in
“C++” has been developed, which is capable of predicting engine dependent parameters for turbojet
engine at varying independent parameters. In conventional configurations, the high temperature and
high pressure gas expand through the turbines which provide just enough power to drive the
compressor and other engine auxiliaries. Further expansion of the gases takes place in the nozzle to
produce the high speed jet. Transpiration cooling technique has been taken for the turbine blade in
the present analysis. Amount of the air bled from the high pressure compressor as a fraction of the
mainstream flow has been calculated. Cooling of turbine blades (first stator only) also examined at
various altitudes and Mach numbers for the various configurations of turbojet and turbofan engines.
Keywords: Gas Turbine, Turbojet Engine, Specific Thrust, Turbine Blade Cooling.
INTERNATIONAL JOURNAL OF ADVANCED RESEARCH IN ENGINEERING
AND TECHNOLOGY (IJARET)
ISSN 0976 - 6480 (Print)
ISSN 0976 - 6499 (Online)
Volume 5, Issue 10, October (2014), pp. 21-33
© IAEME: www.iaeme.com/ IJARET.asp
Journal Impact Factor (2014): 7.8273 (Calculated by GISI)
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IJARET
© I A E M E

22
1. INTRODUCTION
Engines are installed on aircraft for propulsion. Exhaust gases from aircraft engines exit at
the rear of the engine nozzle, causing the aircraft to move forward and air to pass over the aircraft
wings. This air motion creates lift. Most modern aircraft use gas turbine engines to produce the
required thrust force. Such engines are relatively light and compact and have a high power to weight
ratios. Aircraft gas turbines operate in an open cycle based on the Brayton cycle. Actual jet cycles
are not ideal due to thermodynamic losses. The gases in the jet cycle are expanded in the gas turbine
to the pressure that allows generation of the power needed by the compressor, generator, hydraulic
pump, and other work-consuming devices. Other gas turbine-based aerospace engines exist beyond
the turbojet, including turbofan and turboshaft engines, but only turbojets are considered here.
Production of jet aircraft began in 1945, and the technology has advanced significantly since then.
Efforts to increase power and reduce noise and fuel consumption have led to improvements in
turbojet and turbofan engines [1]. Nearly 16,800 jet aircraft were operating globally in 2007, and this
number is expected to increase to 35,300 by 2024 [2]. Due to decreasing world oil reserves,
increasing oil prices and increasing environmental concerns, efforts have increased to improve
system efficiency. Exergy analysis is a practical and useful tool for such activities, with many
engineers and scientists suggesting that exergy analysis is a highly effective method for evaluating
and enhancing thermodynamic performance, and superior to energy analysis [3]. The state of present
technologies in technical and also non-technical practice implies creation of growing complexity of
systems. Turbojet engine as a complex system is multidimensional highly parametric system with
complex dynamics and non-linearities. Its particular property is operation in a broad spectrum of
changes in environment (e.g., temperatures from -60 to +40 °C). If we want to secure optimal
function of such system, it is necessary to develop models of a turbojet engine for the prediction of
the results that would be closed to real system. Furthermore we need to design a control system that
will secure operation converging to optimal one in all eventual states of environment and also inner
states of the system represented by its parameters. Except implementation of classic algorithms of
control, it is possible to design such systems of control and models by use of progressive methods of
artificial intelligence [4]. Measurement programs designed to ascertain the vulnerability and
survivability of gas turbine engines in airborne weapon systems are currently producing operational
results which apply to a broad range of combat scenarios. Battlefield situations have been postulated
in triservice studies [5, 6]. However, the reported work does not consider the response of the engine
to ingestion of dust-laden air. Survivability studies for airborne command aircraft must consider
potential engine degradation caused by the ingestion of atmospheric dust in the operational theater.
The reality of this engine degradation mechanism has been dramatically emphasized by recent
incidents in which gas-turbine-powered air transports have attempted to traverse volcanic ash clouds.
Motivated by the disastrous nature of these encounters on engine performance, the measurement
program described in this paper was initiated. Gas turbine engines are routinely tested for the effect
of ingestion of solid particles according to the procedures of Military Specification MIL-E-5007D,
the commonly known Arizona road-dust test. Within the last several years the material used in the
road-dust test was changed from an earthlike material to a crushed quartz material. Encounters with
volcanic clouds and clouds of earthlike soils are felt to be much different than the MIL-E-5007D test
in material chemical composition, particle size distribution, and cloud concentration [7].
2. MATERIALS AND METHODS
2.1 Modeling of the Components
Early generation jet engines were pure turbojets, designed initially to use a centrifugal
compressor, and very shortly afterwards began to use axial compressors for a smaller diameter to the

23
overall engine housing. They were used because they were able to achieve very high altitudes and
speeds, much higher than propeller engines, because of a better compression ratio and their high
exhaust speed. However, they were not very fuel efficient.
Fig. 1: Schematic of simple turbo jet engine
Fig 1 represents the single-spool (or single-shaft) turbojet engine. In which air enters in an
intake before being compressed to a higher pressure by a compressor. The compressed air passes in
to a combustor, where it is mixed with a fuel and ignited. The hot combustion gases then enter a
windmill-like turbine, where power is extracted to drive the compressor. Although the expansion
process in the turbine reduces the gas pressure (and temperature) somewhat, the remaining energy
and pressure is employed to provide a high-velocity jet by passing the gas through a propelling
nozzle. This process produces a net thrust opposite in direction to that of the jet flow.
2.1.1 Atmospheric Model
As per the ISA, pressure altitude is not set by the elevation of the point above sea level.
Standard day temperature falls at the rate of approximately 60
C per 1000 m [8] until a pressure
altitude of 11000 m which is referred to the tropopause (the region below this is the troposphere,
and that above it the stratosphere) and is given as
(1) If Altitude ሺmሻ ൑ 11000
Tୟ ൌ Tୱ୪ െ 0.00649 ൈ Altitude (1)
pୟ ൌ pୱ୪ ቀ
୘౗
୘౩ౢ
ቁ
ହ.ଶହ଺
(2)
(2) If 11000 ൏ Altitude ሺmሻ ൑ 25000
Tୟ ൌ 216.66 (3)
pୟ ൌ 0.2265. eଵ.଻ଷିሺ଴.଴଴଴ଵହ଻ൈ୅୪୲୧୲୳ୢୣሻ
(4)
(3) If Altitude ሺmሻ ൐ 25000
Tୟ ൌ 141.89 ൅ ሺ0.002991 ൈ Altitudeሻ (5)
pୟ ൌ 0.02488 ቀ
୘౗
ଶଵ଺.଺଺
ቁ
ହ.ଶହ଺
(6)

24
3.1.2 Gas Model
The thermodynamic properties of air and products of combustion are calculated by
considering variation of specific heat and with no dissociation. Specific heats against temperature
variation have been published in many references such as [9, 10, 11].
2.1.2 Diffuser Model
At inlet of the diffuser the condition of the air will be given by,
୘బ,౟
୘౟
ൌ ቂ1 ൅ ቀ
γ౗౟౨ିଵ
ଶ
ቁ . Mଶ
ቃ (7)
୮బ,౟
୮౟
ൌ ቀ
୘బ,౟
୘౟
ቁ
γ౗౟౨
γ౗౟౨షభ
for M ൑ 1 (8)
୮బ,౟
୮౟
ൌ ቀ
୘బ,౟
୘౟
ቁ
γ౗౟౨
γ౗౟౨షభ
ൈ ൬
୮బ,౛
୮బ,౟
൰
ୱ୦୭ୡ୩
for 1 ൏ ‫ܯ‬ ൑ 5 (9)
where ൬
୮బ,౛
୮బ,౟
൰
ୱ୦୭ୡ୩
ൌ 1 െ 0.075ሺM െ 1ሻଵ.ଷହ
(10)
At exit of the diffuser the condition of the air will be given by,
T଴,ୣ ൌ T଴,୧ (11)
୮బ,౛
୮౟
ൌ ቂ1 ൅ ηୢ
. ቀ
γ౗౟౨ିଵ
ଶ
ቁ . Mଶ
ቃ
γ౗౟౨
γ౗౟౨షభ
(12)
2.1.3 Compressor Model
The purpose of a compressor is to increase the total pressure of the gas stream that required
by the cycle by absorbing the minimum possible shaft power.
Mass balance for compressor:
mሶ ୟ୧୰,୧ ൌ mሶ ୟ୧୰,ୣ ൅ ∑ mሶ ୡ୪ (13)
Energy balance for compressor in terms of enthalpy and mass:
Wୡ ൌ mሶ ୟ୧୰,ୣ . hୟ୧୰,ୣ ൅ ∑ mሶ ୡ୪ . hୡ୪ െ mሶ ୟ୧୰,୧. hୟ୧୰,୧ (14)
where mሶ ୡ୪ is the mass of the cooling air which is extracted from the compressor stages. It will be
calculated after turbine cooling flow calculation.
2.1.4 Combustion Chamber Model
The purpose of the combustion system of an aircraft gas turbine engine is to increase the
thermal energy of a flowing gas stream by combustion. Inflowing air is diffused at the entrance of
the burner.
Mass balance for the combustion chamber component will be given by
mሶ ୥ୟୱ ൌ mሶ ୟ୧୰,୧ ൅ mሶ ୤,ୠ (15)

25
Energy balance for the combustion chamber will be as follows
ηୠ
. mሶ ୤,ୠ. LCV୤ ൌ mሶ ୥ୟୱ,ୣ. h୥ୟୱ,ୣ െ mሶ ୟ୧୰,୧. hୟ୧୰,୧ (16)
The pressure at combustor exit is given by
pୠ,ୣ ൌ p ୠ,୧ െ ∆pୠ (17)
where ∆pୠ is the pressure loss in the combustion chamber taken as 2% of pୠ,୧, for
simplicity.
The fuel to air ratio (FAR) is calculated as,
୫ሶ ౜,ౘ
୫ሶ ౗౟౨
ൌ
ୡ౦ౝ౗౩.୘బ,౛ିୡ౦౗౟౨.୘బ,౟
ηౘ.୐େ୚౜ିୡ౦ౝ౗౩.୘బ,౛
(18)
LCV୤ is the lower calorific value of the fuel.
2.1.5 Gas Turbine Model
Turbine produced power to drive engine compressor due to expansion of gas stream. A
general model that represents the gas turbine with turbine blade cooling has been developed. The
model is intended for use in cycle analysis applications.
Work compatibility of turbine and compressor is given as follows:
Wୡ ൅ Wୟ୳୶ ൌ η୫
W୲ (19)
Waux, auxiliary power, required for powering accessories is tapped out from the high pressure turbine.
It is assumed to be 10% of the compressor work. Since
W୲ ൌ
୛ౙ
ηౣ
ൈ 1.10 (20)
Here W୲ is the work developed by high pressure turbine, Wୡ is the work required by the high
pressure compressor and η୫
is the mechanical efficiency which is used to account for the losses due
to windage, bearing friction, and seal drag and is defined as
η୫
ൌ
mechanical power output
mechanical power input
It is to be noticed that the auxiliary work taken should be balanced only once with high pressure
turbine work.
Mass and energy balances yield the turbine work, as given below.
Mass balance for the gas turbine component will be given by
mሶ ୥ୟୱ,୧ ൌ mሶ ୥ୟୱ,ୣ ൅ ∑ mୡ୪ሶ (21)
where

26
mሶ ୥ୟୱ,୧ ൌ mሶ ୟ୧୰,୧ ൅ mሶ ୤,ୠ (22)
and ∑ mୡ୪ሶ represents the sum of air coolant flow rates to the stage (stator and rotor).
Energy balance in terms of enthalpy is given as
mሶ ୥ୟୱ,୧. h୥ୟୱ,୧ ൅ mሶ ୡ୪,୧. hୡ୪,୧ ൌ mሶ ୥ୟୱ,ୣ. h୥ୟୱ,ୣ൅ W୲ (23)
2.1.6 Cooling Model
The cooling of turbine nozzle guide vanes or blades and the rotor is required to limit turbine
blade temperature within the metallurgical limits of the turbine blade material. The cooling air is
drawn off from the high-pressure compressor. It is used to cool the high-pressure and low-pressure
turbine nozzle guide vanes or blades as well as their respective rotors. Due to this air-cooling, there
will be mixing and heat transfer losses in the expansion path and as a result the turbine exhaust
temperature will be affected. The turbine blade cooling model is based on the work of Horlock et al.
[12] and Sanjay et al. [13]. Each stage of turbine is divided into two cooled elements- stator row and
rotor row.
Transpiration cooling is considered in the present work. Fig. 2 shows the simple model for
transpiration cooling.
For an internally cooled turbine configuration, the ratio of coolant to main gas flow rates
൬
୫ሶ ౙౢ
୫ሶ ౝ
൰ required to cool the gas to Tb at the blade surface is proportional to the difference of enthalpy,
which drives the heat transfer from gas to the blades to the ability of the coolant to absorb heat,
which is also termed as cooling factor Rc.
Thus
mሶ cl
mሶ g
=
Heat transfer to blades
Ability of coolant to absorb heat
‫ן‬
hg,iିhbl
hcl,eିhcl,i
(24)
A concept of isothermal effectiveness for transpiration air-cooling (ηiso) trans is introduced and
the resulting cooling factor is given by
ሺRcሻtrans=
൫Tg,iିTbl൯. cp,g. ቀ1-൫ηiso൯
trans
ቁ
ε൫TblିTcl,i൯. cp,cl
(25)
The value of (ηiso)trans= 0.5
For air transpiration cooling, the value of cp,cl will be taken for that air. Thus the cooling
requirement for transpiration cooling is given as follows:
୫ሶ ౙౢ
୫ሶ ౝ
=Sതtin. ቂ
Sg
t cosα
. Fsaቃ . ሾRcሿtrans=0.0156ሾRcሿtrans (26)
The total pressure losses in mixing of coolant and mainstream are expressed as
∆୮
୮
ൌ 0.07
୫ሶ ౙౢ,౟
୫ሶ ౝ,౟
(27)

27
Fig 2: Transpiration air-cooling model for a single row of gas turbine
Mass of coolant required per kg of gas flow is given by following equations
mሶ cl
mሶ g
=0.0156. ሾRcሿtrans (28)
where
ሺRcሻtrans=
൫Tg,i-Tbl൯. cp,g. ቀ1-൫ηiso൯
trans
ቁ
ε൫TblିTcl,i൯. cp,cl
(29)
The total pressure loss in mixing of coolant and mainstream are taken as 2% in this work.
2.1.7 Nozzle Model
The jet nozzle suffers from the aerodynamic losses mainly due to skin friction, which is
modeled by introducing the concept of nozzle efficiency [14]. The nozzle may be choked or
unchoked and may be decided using the critical pressure ratio.
Choked Nozzle and Unchoked Nozzle: If
୮బ,౟
୮౗
൐
୮బ,౟
୮ౙ
, then nozzle is choked. In this condition the
calculation is based on the following equations.
The jet velocity is given by
C୬,ୣ ൌ ටቀγ୥
. R୥. Tୡቁ (30)
The mass flow rate is expressed by
m୥,ୣሶ ൌ ρ୬,ୣ
. A୬,ୣ. C୬,ୣ (31)
If
୮బ౟
୮౗
൏
୮బ౟
୮ౙ
the nozzle is unchoked and the jet velocity is given by
C୬,ୣ ൌ ටൣ2c୮୥. ൫T଴,୧ െ T଴,ୣ൯൧ (32)
2.1.8 Governing Equations for performance prediction and software developments
Following equations are used to calculate specific thrust and thrust specific fuel consumption
of a turbojet engine. The cycle performance and optimization are achieved through software

28
developed in “C++” language and then graphs plotted in the menu driven software “origin 61”.
Under different nozzle conditions it is expressed as given below:
If nozzle is chocked, the net specific thrust comprise by the two components, namely
Momentum thrust given by
F୫ ൌ mሶ ୥. ሺC୬,ୣ െ Cୟሻ (33)
Pressure thrust given by
F୔ ൌ A୬,ୣ. ൫p଴,ୣ െ pୟ൯ (34)
Therefore the net thrust Fୱ is given by the following equation
Fୱ ൌ F୫ ൅ F୔ (35)
If nozzle is unchoked, net specific thrust is directly expressed as
Fୱ ൌ mሶ ୥. C୬,ୣ (36)
Thrust Specific Fuel Consumption is expressed as
TSFC ൌ
୫ሶ ౜
୊౩
(37)
3. RESULTS AND DISCUSSIONS
Table 1: Input values for the present analysis
S. No. Symbols Values
1 TA (Atmospheric Temperature) 288.15 K
2 PA (Atmospheric Pressure) 1.01325 bar
3 R (Gas constant for air) 287 kJ/kgK
4 ETAD (Diffuser Efficiency) 0.95
5 ETAF (Fan Efficiency) 0.87
6 ETAC (Compressor Efficiency) 0.87
7 ETAP (Turbine Efficiency) 0.87
8 RG (Gas constant for gas) 0.296 J/kgK
9 ETAM (Mechanical Efficiency) 0.99
10 TRS (Maximum number of stages) 25
11 DOR (Degree of reaction) 0.5
12 C (cooling factor) 0.03
13 ETACL (Cooling Efficiency) 0.8
14 TBL (Blade Temperature) 1122 K
15 CCL (Cooling constant) 2
16 MA (Mass of air rate flow) 1 kg/s
17 ETAB (Blade efficiency) 0.98
18 ETAJ (Nozzle Efficiency) 0.97
19 CV (Calorific Value of Fuel) 43124 kJ/kg
20 ∆p (Pressure drop in Combustion chamber) 2%
21 M (Mach Number) 0.9 - 1.5
22 OPR (Overall Compression Pressure Ratio) 10 – 30
23 TIT (Turbine Inlet Temperature) 1000 – 1500K

29
In the following sections the results of the present work have been discussed in the form of
graphs plotted in the menu driven software “Origin 51”. All the results are shown for various
parameters of Altitudes, Turbine inlet temperature and overall pressure ratio. The variations of
Specific thrust, Thrust specific fuel consumption and mass of fuel required in the combustion
chamber have been plotted with respect to the turbine inlet temperatures.
1000 1100 1200 1300 1400 1500
150
200
250
300
350
400
450
500
550
600
650
700
750
TurbineInlet Temperature(K)
M=0.9
Alt.=10000m
OPR=10
OPR=15
OPR=20
OPR=25
OPR=30
SpecificThrust(N)
Figure 3: Variations of Specific thrust with TIT
Figure 3 represents the variation of Specific thrust with turbine inlet temperature for a given
Mach number and Altitude for various overall pressure ratios. On increasing the overall pressure
ratio of the compressor, the specific thrust increases for a given value of TIT.
1000 1100 1200 1300 1400 1500
0.010
0.015
0.020
0.025
0.030
0.035
0.040
0.045
M=0.9
Alt.=10000 m
MassofFuelRequiredinCombustionChamber(kg/s)
Turbine Inlet Temperature (K)
OPR=10
OPR=15
OPR=20
OPR=25
OPR=30
Figure 4: Variations of Mass of fuel required in combustion chamber with TIT

30
1000 1100 1200 1300 1400 1500
0.15
0.16
0.17
0.18
0.19
0.20
0.21
0.22
TurbineInlet Temperature(K)
M=0.9
Alt.=10000m
OPR=10
OPR=15
OPR=20
OPR=25
OPR=30
ThrustSpecificFuelConsumption(kg/hN)
Figure 5: Variations of Thrust specific fuel consumption with TIT
Figure 4 represents the variations of mass of fuel required in the combustion chamber with
TIT for a given low Mach number and low altitude. On increasing the pressure ratio the fuel required
in the combustion chamber decreases due to increase in temperature of inlet air before entering into
the combustion chamber.
Figure 5 shows the variations of Thrust specific fuel consumption with TIT for a given Mach
number and Altitude but different pressure ratio. The thrust specific fuel consumption decreases on
increasing the TIT in the low range say 1000 to 1200 K. After that range of temperature, the thrust
specific fuel consumption increases on increasing the TIT for a given pressure ratio.
1000 1100 1200 1300 1400 1500
0.000
0.005
0.010
0.015
0.020
Massofcoolantperkgofmassflowofair
M=0.9
Alt.=10000 m
mc
/mg
for 1st Stator
mc
/mg
for 1st Rotor
mc
/mg
for 2nd Stator
mc
/mg
for 2nd Rotor
Figure 6: Variations of Mass of coolant required with TIT
Figure 6 represents the variations of Mass of coolant required per kg of air flow with TIT. It
is clearly shown from the figure that the turbine blades are safe up to temperature 1123 K. When the
temperature reached above that value, the coolant required to cool the turbine blade and increases the
life.
Figure 7 represents the variations of Specific thrust with TIT for various Mach number. At
higher Mach number or for sonic velocity, the specific thrust is higher as compares to the subsonic
velocity but in the higher turbine inlet temperature range.

31
1000 1100 1200 1300 1400 1500
250
300
350
400
450
500
550
600
650
700
750
OPR=20
Alt.=10000 m
SpecificThrust(N)
M=0.9
M=1.1
M=1.3
Figure 7: Variations of Specific thrust with TIT
1000 1100 1200 1300 1400 1500
0.010
0.015
0.020
0.025
0.030
0.035
0.040
OPR=20
Alt.=10000 m
Massoffuelrequiredincombustionchamber
M=0.9
M=1.1
M=1.3
Figure 8: Variations of Mass of fuel required in combustion chamber with TIT
Figure 8 represents the variations of mass of fuel required in the combustion chamber with
TIT for different Mach numbers and low altitude. On increasing the mach number the fuel required
in the combustion chamber decreases.
1000 1100 1200 1300 1400 1500
0.130
0.135
0.140
0.145
0.150
0.155
0.160
0.165
0.170
0.175
0.180
0.185
0.190
0.195
0.200
OPR=20
Alt.=10000 m
ThrustSpecificFuelConsumption(kg/hN)
M=0.9
M=1.1
M=1.3
Figure 9: Variations of Thrust Specific Fuel Consumption with TIT

32
Figure 9 shows the variations of Thrust specific fuel consumption with TIT for different
Mach number and a given Altitude and pressure ratio. The thrust specific fuel consumption decreases
on increasing the Mach number.
4. CONCLUSION
A parametric thermodynamic study has been carried out for a single spool turbojet engine
presenting a summary of point design performance estimating using simplified component
technology assumptions. The present work may be concluded as follows-
1) The results obtained for single spool turbojet engine shows high specific thrust produces at
higher Mach number for a given Turbine Inlet Temperature.
2) Specific thrust is higher for a given Mach number and turbine inlet temperature on increasing
the pressure ratio of the compressor.
3) Mass of fuel required in the combustion chamber decreases on increasing he pressure ratio
for a given value of TIT.
4) At any altitude the TSFC first decreases at lower TIT and then increases with increase in TIT
and Mach number, whereas reverse is true with increase in altitude at any Mach number.
5) Mass of cooling air required will depend on the TIT. The cooling required after a certain
temperature of gases entering into the turbine blade. The cooling is necessary when the
temperature of the gases reached above metallurgical limit of the turbine blade, therefore it is
important to use cooling air above 1123 K to increases the life of the turbine blade.
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Thesis, Motilal Nehru National Institute of Technology, Allahabad.
[10] Fedrik Hanglind, and Riti Singh (2006), “Design of aero gas turbines using hydrogen”,
Transaction of ASME Vol. 128, pp. 754-764.

33
[11] Lakshminarayana, B. L. (1996), “Fluid Dynamics and Heat Transfer in Turbomachinery”,
John Wiley and Sons, New York, Chapter 7.
[12] Horlock, J. H. (2001), “The Basic Thermodynamics of Turbine Cooling”, Journal of
Turbomachinery, Vol.123, pp. 583-591.
[13] Sanjay, Onkar Singh, and Prasad, B. N. (2008), “Comparative performance analysis of
cogeneration gas turbine cycle for different blade cooling means”, International Journal of
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[14] Philip P. Walsh and Paul Fletcher, “Gas Turbine Performance”, Second Edition 2004 by
Blackwell Science Ltd a Blackwell Publishing company Editorial Offices: Blackwell Science
Ltd, 9600 Garsington Road, Oxford OX4 2DQ, UK.
[15] P.S. Jeyalaxmi and Dr.G.Kalivarathan, “CFD Analysis of Turbulence in a Gas Turbine
Combustor with Reference to the Context of Exit Phenomenon”, International Journal of
Advanced Research in Engineering & Technology (IJARET), Volume 4, Issue 2, 2013,
pp. 1 - 7, ISSN Print: 0976-6480, ISSN Online: 0976-6499.
[16] Yousif Khudhair Abbas, “Effects of Sunlight Intensity on Turbo Jet Engine of Aircraft”,
International Journal of Mechanical Engineering & Technology (IJMET), Volume 5, Issue 3,
2014, pp. 31 - 39, ISSN Print: 0976 – 6340, ISSN Online: 0976 – 6359.
[17] Prabhat Kumar Sinha, Mohdkaleem, Ivan Sunit Rout and Raisul Islam, “Analysis &
Modelling of Thermal Mechanical Fatigue Crack Propagation of Turbine Blade Structure”,
International Journal of Mechanical Engineering & Technology (IJMET), Volume 4, Issue 3,
2013, pp. 155 - 176, ISSN Print: 0976 – 6340, ISSN Online: 0976 – 6359.
[18] Aram Mohammed Ahmed and Dr. Mohammad Tariq, “Thermal Analysis of a Gas Turbine
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AUTHOR’S DETAIL
I am Rana Adil Abdul-Nabe, I am working as a Teacher, Ministry of Higher Education, Institute of
the Musaib Technical Institute, I have completed my Bachelor at Musaib Technical College,
Babylon, Republic of Iraq. I am going to complete M. Tech. in Mechanical Engineering (Thermal
Engineering) in SSET, Faculty of Engineering, Sam Higginbottom Institute of Agriculture,
Technology & Sciences, Allahabad, India.
I am Dr. Mohammad Tariq working as an Assistant Professor (Senior) in the
Department of Mechanical Engineering, Shepherd School of Engineering and
Technology, SHIATS-DU, Allahabad. I have published more than 35 research
papers in National/International Journals. I have guided more than 25 students for
M. Tech. so far.

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